A gas turbine engine generally includes a fan and a core arranged in flow communication with one another. Additionally, the core of the gas turbine engine generally includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. In operation, air is provided from the fan to an inlet of the compressor section where one or more axial compressors progressively compress the air until it reaches the combustion section. Fuel is mixed with the compressed air and burned within the combustion section to provide combustion gases. The combustion gases are routed from the combustion section to the turbine section. The flow of combustion gases through the turbine section drives the turbine section and is then routed through the exhaust section, e.g., to atmosphere.
In general, turbine performance and efficiency may be improved by increased combustion gas temperatures. Non-traditional high temperature materials, such as ceramic matrix composite (CMC) materials, are more commonly being used for various components within gas turbine engines. For example, because CMC materials can withstand relatively extreme temperatures, there is particular interest in replacing components within the flow path of the combustion gases with CMC materials. However, even though CMC components may withstand more extreme temperatures than typical components, CMC components still may require cooling features or reduced exposure to the combustion gases to decrease a likelihood of negative impacts of increased combustion gas temperatures, e.g., material failures or the like.
More specifically, CMC airfoils for gas turbine engines typically have a cavity for receipt of a cooling fluid located near a forward end of the airfoil, i.e., proximate a leading edge of the airfoil. Often, an aft end of the airfoil, i.e., proximate a trailing edge of the airfoil, does not have a cavity or other feature for receipt of a cooling fluid and thus remains uncooled, which can produce a large temperature gradient between the forward end and the aft end of the airfoil. A large temperature gradient across the airfoil can increase the thermal stress on the airfoil, which can lead to material failures or other negative impacts on turbine performance.
Therefore, improved cooling features for CMC components that overcome one or more disadvantages of existing components would be desirable. In particular, a turbine nozzle segment for a gas turbine engine having cooling features in a trailing edge portion of a CMC airfoil of the turbine nozzle segment would be beneficial. Moreover, a turbine nozzle segment for a gas turbine engine having cooling features in a trailing edge portion of a CMC airfoil of the turbine nozzle segment that even out cooling of the airfoil would be desirable. Methods of cooling an airfoil of a turbine nozzle segment by supplying cooling fluid from a cavity defined by an inner and/or outer band of the turbine nozzle segment directly to an internal cooling passage defined in a trailing edge portion of the airfoil would be advantageous.